The present invention relates to a low-thrust cryogenic propulsion module applicable to a conventional launcher or to a recoverable launcher.
The invention relates more particularly to a low-thrust cryogenic propulsion module for transferring the orbit of a satellite, the propulsion module being suitable for being integrated in the satellite or for constituting a separate propulsion stage.
A major preoccupation in the field of launching satellites lies in improving the mass injected into geostationary orbit for given launcher mass.
The most conventional method consists in injecting a satellite into a geostationary transfer orbit (GTO) and then in transferring the satellite into geostationary orbit using a two-liquid apogee engine, the two-liquid propulsion module being integrated in the satellite.
Proposals have also been made to transfer a satellite from a low orbit into a geostationary orbit by means of a solar thermal stage using liquid hydrogen.
Such a method is described, for example, in the article by J. A. Bonometti and C. W. Hawk entitled xe2x80x9cSolar thermal rocket research apparatus and proposed testingxe2x80x9d (University of Alabama, 1994).
That method is also mentioned in the article by J. M. Shoji published in Progress in Astronautics and Aeronautics, AIAA, Vol. 87, (pp. 30 to 47), and entitled xe2x80x9cPotential of advanced solar thermal propulsion. Orbit raising and maneuvering propulsion: research status and needsxe2x80x9d.
In that known method, which is shown in FIG. 2, light from the sun is concentrated by a parabolic mirror 5 onto a solar furnace 6 in which hydrogen is heated to a temperature of about 2000 K. The hydrogen is then expanded in a nozzle of a thruster 8 to deliver a high ejection speed (7500 meters per second (m/s) to 8000 m/s), giving a specific impulse of about 750 seconds (s) to 800 s. FIG. 2 is a diagram showing such an arrangement with a satellite 2 connected firstly to a launcher via an interface 1 and secondly to a hydrogen tank 3 via a truss 4. Reference 7 represents diagrammatically a device for acquiring liquid hydrogen in order to feed the solar furnace 6 and the thruster 8.
Such a device, which has never been used in practice, ought theoretically to make it possible to increase the mass that it injected into geostationary orbit. Nevertheless, that configuration presents various drawbacks.
In particular, in order to reach a temperature of 2000 K, it is necessary to use a solar flux concentration factor of 5000 to 8000, which requires a mirror of very good quality, which is very difficult to obtain when subject to constraints limiting on-board mass. In addition, pointing towards the sun must be very accurate, of the order of xc2x15 minutes of arc about two axes, which gives rise to problems in attitude control.
The size of the liquid hydrogen tank 3 also constitutes a difficulty. For example, in order to obtain total delivered impulse of 30 meganewton-seconds (MN.s), it is necessary to use a tank containing 4000 kilograms (kg) of liquid hydrogen which thus presents a volume of 60 cubic meters (m3) (which, for example, implies a diameter of 4.2 m and a height of 5 m).
Developments in orbit transfer systems based on a solar thermal stage are in serious difficulty due to those drawbacks.
In another technique for increasing the mass placed in geostationary orbit, use is made of a launcher top stage of the cryogenic type that makes it possible to use tanks of relatively small volume that are easier to integrate in the launcher. Thus, to obtain a total impulse of 30 MN.s, a liquid hydrogen and liquid oxygen cryogenic stage requires a propellant mass of 6600 kg, but the total volume of the tanks is only 22 m3.
Cryogenic stages currently in use nevertheless require turbopumps to be used, and that increases their cost.
Certain authors have proposed making cryogenic stages that are fed by means of pressure, without using turbopumps, but those concepts have not given rise to concrete implementations. In practice, the hydrogen must always be at a higher pressure than the oxygen in order to perform regenerative cooling of the combustion chamber. It follows that the mass of helium required for pressurization purposes becomes prohibitive.
The invention seeks to remedy the above-mentioned drawbacks and in particular to enable the orbit of a satellite to be transferred using a device that is simpler, lighter, and more compact than prior art devices while avoiding the use of turbopumps and making it possible to use both thrusters and propellant tanks of reasonable size reducing the bulk of the propulsion stage in question required for transferring the orbit of the satellite.
These objects are achieved by a low-thrust cryogenic propulsion module presenting thrust lying in the range 100 N to 1000 N, the module being characterized in that it comprises at least one main cryogenic thruster whose combustion pressure lies in the range 2 bars to 10 bars, at least two attitude-controlling secondary thrusters, at least first and second feed tanks for feeding cryogenic propellants, means for intermittently pressurizing said feed tanks, and means for triggering intermittent firing of the main cryogenic thruster during intermittent pressurization of said feed tanks, the duration between two successive firings lying in the range about 1 hour (h) 30 minutes (min) to 12 h, in that the means for intermittently pressurizing a feed tank comprises at least one heat exchange circuit associated with a heat accumulator and with means for circulating a predetermined quantity of a propellant through said heat exchanger, and in that it further comprises means for heating the heat accumulator in the periods that lie between two consecutive firings.
The heat accumulator associated with the propellant tank can be heated, at least in part, by means of a solar collector, e.g. using a plane solar collector having an absorptance/emissivity ratio (xcex1/xcex5) greater than one, and which is provided with superinsulation on its rear face.
Nevertheless, the heat accumulator can also be heated at least in part by recovering heat losses from a fuel cell operating by means of evaporated propellants.
The fuel cell can be fed with cold propellant vapor coming from a heat exchanger for keeping the temperature at which propellant is taken from a propellant tank constant.
The heat accumulator can also be heated, at least in part, by electrical heating.
Heat accumulation within the heat accumulator is advantageously performed by a material that changes phase, such as an alkali metal or a hydrocarbon.
In a particular embodiment, the cryogenic propulsion module comprises first and second propellant tanks for feeding the main thruster and the propellants are fully vaporized in the heat accumulators associated with the tanks so as to guarantee a constant mixture ratio.
In an advantageous embodiment, the cryogenic propulsion module has at least first and second main propellant tanks and at least first and second secondary propellant tanks constituting buffer tanks, which secondary tanks can be pressurized by said pressurizing means and are dimensioned in such a manner as to enable orbital maneuvering to be performed while feeding the main thruster intermittently and so as to be completely emptied at the end of firing, means being provided for re-feeding said secondary tanks from the corresponding main tanks between two successive firings, with the pressure of the main tanks being kept below the pressure at which the main thruster is fed.
Under such circumstances, in a particular embodiment, a secondary tank is covered in thermal insulation and is mounted inside a main tank.